DNS 1-6 Description: Difference between revisions
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==Boundary conditions | TO UPDATE !!!== | ==Boundary conditions | TO UPDATE !!!== | ||
Although for DNS the inflow boundary conditions are different than RANS, to allow a valid comparison they must guarantee the same boundary layer properties at a given point, hereinafter referred to as checkpoint, upstream of the rounded step, i.e., <math>{x_{ckp}/H = -3.5}</math>. | |||
At this position <math>{x_{ckp}}</math>, the properties to be matched are the Reynolds number based on the momentum thickness <math>{Re_{\theta} = 1,780}</math>, the Reynolds number based on the friction velocity <math>{Re_{\tau} = 700}</math> and the boundary layer thickness <math>\delta_{99}/H=0.241</math>. | |||
As a technique to promote the laminar-turbulent transition of the boundary layer and reduce the upstream length of the domain, a local tripping term inspired by the work of [[lib:DNS_1-5_description#5|Housseini ''et al.'' (2016)]] and [[lib:DNS_1-5_description#6|Schlatter and Örlü (2012)]] is used. | |||
To define the mesh density and the computational domain size that ensure the target boundary layer integral parameters at <math>{x_{chp}}</math>, a precursory computational campaign for the turbulent flow over a flat plate was performed. | |||
According to the outcomes of this campaign, the inlet boundary is located at <math>{x/H = -12.7}</math>, where the Blasius laminar velocity profile computed at <math>{Re_x=650,000}</math>, the uniform static pressure <math>{P_{s,ref}}</math> and the uniform total temperature <math>{T_{t,inflow}}</math> are imposed, see [[Lib:UFR_3-36_Test_Case#table2|UFR 3-36: Table 2]] for dimensional values of each quantity. At location <math>{x/H = -12.1}</math>, within the laminar boundary layer region, the tripping source term is activated to promote the transition to turbulence, see [[Lib:DNS_1-5_description#figure2|Fig. 2]]. | |||
At the outlet boundary, placed at <math> x/H=24.0 </math>, the static pressure <math>{P_{s,ref}}</math> is imposed with an exit-pressure outflow boundary condition. To mitigate spurious perturbations possibly originating at the outlet boundary, the mesh is coarsened in the streamwise direction for <math>{x/H > 13.82}</math>. The upper boundary is a permeable far-field Riemann boundary condition located <math>180.0 H</math> from the no-slip adiabatic wall downstream the smooth step and computed via the exact Riemann solver. Finally, side planes are considered as periodic with a distance from each other of <math>{\Delta z = 3 H}</math>. | |||
For this particular flow configuration, we aim at simulating a transitional boundary layer in order to replicate the flow conditions of the experiment. To do so, a Blasius velocity profile is imposed at the inlet corresponding to <math>Re_x=900,000</math> with a uniform temperature. A laminar boundary layer is then established, and a flow perturbation is introduced at <math>Re_x=950,000</math> with amplitude <math>A_T=\frac{\rho_{ref} U_{ref}^2}{T}</math> to trigger transition to turbulence. The total length of the boundary layer was determined such that the turbulent boundary layer thickness upwind the airfoil reaches half of the experimental value. Symmetry conditions are imposed at the lateral boundary conditions (<math>z=\pm 4T</math>). The bottom boundary is a no-slip adiabatic wall type for <math>xz</math> planes at <math>y=0</math> between <math>x=-12.75T</math> and <math>x=17T</math>, and symmetry type between <math>x=17T</math> and <math>x=67T</math>. The outlet is located away from the profile at <math>x=67T</math>, and the zone between <math>x=17T</math> and <math>x=67T</math> acts as a sponge layer, featuring increasingly coarse elements such that a constant flow field is recovered when reaching the outlet. | For this particular flow configuration, we aim at simulating a transitional boundary layer in order to replicate the flow conditions of the experiment. To do so, a Blasius velocity profile is imposed at the inlet corresponding to <math>Re_x=900,000</math> with a uniform temperature. A laminar boundary layer is then established, and a flow perturbation is introduced at <math>Re_x=950,000</math> with amplitude <math>A_T=\frac{\rho_{ref} U_{ref}^2}{T}</math> to trigger transition to turbulence. The total length of the boundary layer was determined such that the turbulent boundary layer thickness upwind the airfoil reaches half of the experimental value. Symmetry conditions are imposed at the lateral boundary conditions (<math>z=\pm 4T</math>). The bottom boundary is a no-slip adiabatic wall type for <math>xz</math> planes at <math>y=0</math> between <math>x=-12.75T</math> and <math>x=17T</math>, and symmetry type between <math>x=17T</math> and <math>x=67T</math>. The outlet is located away from the profile at <math>x=67T</math>, and the zone between <math>x=17T</math> and <math>x=67T</math> acts as a sponge layer, featuring increasingly coarse elements such that a constant flow field is recovered when reaching the outlet. | ||
Revision as of 15:09, 16 February 2023
Introduction
This test case features a 3:2 semi-elliptic nose with a NACA0020 tail profile mounted on a flat plate, which is representative of the wing-body junction flow problems encountered in applications of aeronautical interest. The flow features the interaction between the incipient turbulent boundary layer and the mounted airfoil and the main physical phenomenon of interest is the horseshoe vortex developping at the junction and the corner separation. This flow is also highly 3D and anisotropic regarding the turbulent stresses. Establishing a DNS database of this flow is of crucial interest since it has been shown that RANS models (both Boussinesq and Reynolds stresses-based models) display strong difficulties in recovering data from the available experiments. Such a database allows for a more thorough availability of the flow field with respect to the experiments and gives the possibility of using Machine Learning or data-assimilation techniques to improve standard RANS models.
Review of previous studies and choice of test case
A thorough listing of existing experimental and numerical studies regarding wing-body junction flows can be found in Gand et al. (2010). The present DNS is based on the configuration considered in the simulations by Apsley & Leschziner (2001), who were based themselves on the experimental studies by Devenport and Simpson (1990) and Fleming et al. (1995). The Reynolds number based on the airfoil thickness is similar to the experiment and its value is 115,000 and the flow is almost incompressible with a Mach number based on the freestream velocity of 0.078. The uDNS setup reproduces the experimental conditions but with half the experimental impacting boundary layer thickness. Therefore a direct comparison with the experimental results is not possible for the present uDNS.
Description of the test case
Geometry and flow parameters
Fig.1 displays a view of the computational domain and flow configuration geometry. The reference length scale is the wing chord , with a corresponding Reynolds number and the chord-to-tickness ratio is 4.254. The computational domain size in the streamwise direction is in the spanwise direction and in the wall-normal direction. The coordinates origin is located at the root leading edge of the airfoil. The is no flow incidence relatively to the wing, corresponding to an angle of attack of 0 degrees.
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| Figure 2: Set-up for the DNS simulation |
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Boundary conditions | TO UPDATE !!!
Although for DNS the inflow boundary conditions are different than RANS, to allow a valid comparison they must guarantee the same boundary layer properties at a given point, hereinafter referred to as checkpoint, upstream of the rounded step, i.e., . At this position , the properties to be matched are the Reynolds number based on the momentum thickness , the Reynolds number based on the friction velocity and the boundary layer thickness . As a technique to promote the laminar-turbulent transition of the boundary layer and reduce the upstream length of the domain, a local tripping term inspired by the work of Housseini et al. (2016) and Schlatter and Örlü (2012) is used. To define the mesh density and the computational domain size that ensure the target boundary layer integral parameters at , a precursory computational campaign for the turbulent flow over a flat plate was performed. According to the outcomes of this campaign, the inlet boundary is located at , where the Blasius laminar velocity profile computed at , the uniform static pressure and the uniform total temperature are imposed, see UFR 3-36: Table 2 for dimensional values of each quantity. At location , within the laminar boundary layer region, the tripping source term is activated to promote the transition to turbulence, see Fig. 2. At the outlet boundary, placed at , the static pressure is imposed with an exit-pressure outflow boundary condition. To mitigate spurious perturbations possibly originating at the outlet boundary, the mesh is coarsened in the streamwise direction for . The upper boundary is a permeable far-field Riemann boundary condition located from the no-slip adiabatic wall downstream the smooth step and computed via the exact Riemann solver. Finally, side planes are considered as periodic with a distance from each other of .
For this particular flow configuration, we aim at simulating a transitional boundary layer in order to replicate the flow conditions of the experiment. To do so, a Blasius velocity profile is imposed at the inlet corresponding to with a uniform temperature. A laminar boundary layer is then established, and a flow perturbation is introduced at with amplitude to trigger transition to turbulence. The total length of the boundary layer was determined such that the turbulent boundary layer thickness upwind the airfoil reaches half of the experimental value. Symmetry conditions are imposed at the lateral boundary conditions (). The bottom boundary is a no-slip adiabatic wall type for planes at between and , and symmetry type between and . The outlet is located away from the profile at , and the zone between and acts as a sponge layer, featuring increasingly coarse elements such that a constant flow field is recovered when reaching the outlet.
References
- Apsley, D.D. and Leschziner, M. (2001): Investigation of Advanced Turbulence Models for the Flow in a Generic Wing-Body Junction. Flow, Turbulence and Combustion, Vol. 67, pp. 25–55
- Gand, F., Deck, S., Brunet, V., and Sagaut, P. (2010): Flow dynamics past a simplified wing body junction. Physics of Fluids, Vol. 22, 115111
- Devenport W.J. and Simpson R.L. (1990): Time-dependent and time-averaged turbulence structure near the nose of a wing-body junction. Journal Fluid Mechanics, Vol. 67, pp. 23–55
- Fleming, J.L., Simpson, R.L., Cowling, J.E. and Devenport, W.J. (1993): An experimental study of wing-body junction and wake flow. Exp. Fluids Vol. 14, pp. 366–378
Contributed by: Francesco Bassi (UNIBG), Alessandro Colombo (UNIBG), Francesco Carlo Massa (UNIBG), Michael Leschziner (ICL/ERCOFTAC), Jean-Baptiste Chapelier (ONERA) — University of Bergamo (UNIBG), ICL (Imperial College London), ONERA
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